US8549834B2 - Gas turbine engine with variable area fan nozzle - Google Patents
Gas turbine engine with variable area fan nozzle Download PDFInfo
- Publication number
- US8549834B2 US8549834B2 US12/909,793 US90979310A US8549834B2 US 8549834 B2 US8549834 B2 US 8549834B2 US 90979310 A US90979310 A US 90979310A US 8549834 B2 US8549834 B2 US 8549834B2
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- US
- United States
- Prior art keywords
- fan nacelle
- fan
- nacelle section
- trailing edge
- section
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
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Classifications
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K1/00—Plants characterised by the form or arrangement of the jet pipe or nozzle; Jet pipes or nozzles peculiar thereto
- F02K1/06—Varying effective area of jet pipe or nozzle
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K1/00—Plants characterised by the form or arrangement of the jet pipe or nozzle; Jet pipes or nozzles peculiar thereto
- F02K1/06—Varying effective area of jet pipe or nozzle
- F02K1/09—Varying effective area of jet pipe or nozzle by axially moving an external member, e.g. a shroud
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K1/00—Plants characterised by the form or arrangement of the jet pipe or nozzle; Jet pipes or nozzles peculiar thereto
- F02K1/38—Introducing air inside the jet
- F02K1/383—Introducing air inside the jet with retractable elements
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K1/00—Plants characterised by the form or arrangement of the jet pipe or nozzle; Jet pipes or nozzles peculiar thereto
- F02K1/38—Introducing air inside the jet
- F02K1/386—Introducing air inside the jet mixing devices in the jet pipe, e.g. for mixing primary and secondary flow
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K1/00—Plants characterised by the form or arrangement of the jet pipe or nozzle; Jet pipes or nozzles peculiar thereto
- F02K1/46—Nozzles having means for adding air to the jet or for augmenting the mixing region between the jet and the ambient air, e.g. for silencing
- F02K1/48—Corrugated nozzles
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K1/00—Plants characterised by the form or arrangement of the jet pipe or nozzle; Jet pipes or nozzles peculiar thereto
- F02K1/54—Nozzles having means for reversing jet thrust
- F02K1/76—Control or regulation of thrust reversers
- F02K1/763—Control or regulation of thrust reversers with actuating systems or actuating devices; Arrangement of actuators for thrust reversers
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K3/00—Plants including a gas turbine driving a compressor or a ducted fan
- F02K3/02—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber
- F02K3/04—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type
- F02K3/06—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type with front fan
Definitions
- the present disclosure relates to a gas turbine engine, and more particularly to a turbofan engine having a variable area fan nozzle (VAFN).
- VAFN variable area fan nozzle
- VAFN variable area fan nozzle
- a nacelle assembly for a bypass gas turbine engine includes a variable area fan nozzle having a first fan nacelle section and a second fan nacelle section.
- the variable area fan nozzle is in communication with a fan bypass flow path, the first fan nacelle section defines an intermittent trailing edge which defines a multiple of ports and the second fan nacelle section defines a multiple of doors, each of the multiple of doors match each of the multiple of ports such that a fan nacelle trailing edge is continuous when the second fan nacelle section is selectively translated to a closed position relative to the first fan nacelle section.
- a gas turbine engine includes a core engine defined about an axis.
- a core nacelle defined at least partially about the core engine.
- a fan nacelle is mounted at least partially around the core nacelle to define a fan bypass flow path and a variable area fan nozzle having a first fan nacelle section and a second fan nacelle section.
- variable area fan nozzle is in communication with a fan bypass flow path
- the first fan nacelle section defines an intermittent trailing edge which defines a multiple of ports
- the second fan nacelle section defines a multiple of doors, each of the multiple of doors match each of the multiple of ports such that a fan nacelle trailing edge is continuous when the second fan nacelle section is selectively translated to a closed position relative to the first fan nacelle section.
- a method of varying a nozzle of a gas turbine engine includes selective translating a second fan nacelle section that defines a multiple of doors relative a first fan nacelle section having a multiple of ports such that a fan nacelle trailing edge is continuous when the second fan nacelle section is selectively translated to a closed position and is intermittent when selectively translated to an open position.
- FIG. 1 is a general schematic partial fragmentary view of an exemplary gas turbine engine embodiment for use with the present invention
- FIG. 2 is a rear view of the engine
- FIG. 3A is a rear view of the engine with the VAFN in a closed position
- FIG. 3B is a rear view of the engine with the VAFN in an open position
- FIG. 4 is a perspective view of one non-limiting embodiment of a VAFN fan nacelle section
- FIG. 5 is a perspective view of another non-limiting embodiment of a VAFN fan nacelle section
- FIG. 6A is a perspective view of the VAFN in a closed position
- FIG. 6B is a perspective view of the VAFN in an open position
- FIG. 7 is a schematic view which illustrates a modulation in the fan nozzle exit area provided by the VAFN;
- FIG. 8 is a sectional view of an interface between the second fan nacelle section and the first fan nacelle section of the VAFN;
- FIG. 9 is an inner perspective view of one non-limiting embodiment of an actuator system for the VAFN.
- FIG. 10 is a sectional view of the actuator system taken along line 10 - 10 in FIG. 9 ;
- FIG. 11 is an inner perspective view of another non-limiting embodiment of an actuator system for the VAFN.
- FIG. 12 is a perspective view of the VAFN fan nacelle section in FIG. 11 ;
- FIG. 13 is an inner perspective view of another non-limiting embodiment of an actuator system for the VAFN.
- FIG. 14 is an expanded perspective view of the VAFN fan nacelle section in FIG. 13 ;
- FIG. 15 is a sectional view of the actuator system taken along line 15 - 15 in FIG. 14 ;
- FIG. 16 is a sectional view of the actuator system taken along line 16 - 16 in FIG. 14 .
- FIG. 1 illustrates a general partial fragmentary schematic view of a gas turbofan engine 10 suspended from an engine pylon P within an engine nacelle assembly N.
- the turbofan engine 10 includes a core engine within a core nacelle 12 that houses a low spool 14 and high spool 24 .
- the low spool 14 includes a low pressure compressor 16 and low pressure turbine 18 .
- the low spool 14 also drives a fan section 20 through a geared architecture 22 .
- the high spool 24 includes a high pressure compressor 26 and high pressure turbine 28 .
- a combustor 30 is arranged between the high pressure compressor 26 and high pressure turbine 28 .
- the low and high spools 14 , 24 rotate about an engine axis of rotation A.
- the engine 10 in one non-limiting embodiment is a bypass geared architecture aircraft engine with a high bypass ratio and a turbofan diameter significantly larger than that of the low pressure compressor 16 .
- the geared architecture 22 may be an epicycle gear train such as a planetary gear system or other gear system with a gear reduction ratio of greater than 2.5:1. It should be understood, however, that the above parameters are only exemplary of one non-limiting embodiment of a geared architecture engine and that this disclosure is applicable to other gas turbine engines including direct drive turbofans.
- a portion of airflow referred to as core airflow, communicates into the core nacelle 12 .
- Core airflow compressed by the low pressure compressor 16 and the high pressure compressor 26 is mixed with the fuel in the combustor 30 and expanded over the high pressure turbine 28 and low pressure turbine 18 .
- the turbines 28 , 18 are coupled for rotation with respective spools 24 , 14 to rotationally drive the compressors 26 , 16 and through the gear train 22 , the fan section 20 in response to the expansion.
- a core engine exhaust E exits the core nacelle 12 through a core nozzle 43 defined between the core nacelle 12 and a tail cone 32 .
- the core nacelle 12 is supported within the fan nacelle 34 by circumferentially spaced structures 36 often referred to as Fan Exit Guide Vanes (FEGVs).
- a bypass flow path 40 is defined between the core nacelle 12 and the fan nacelle 34 .
- the engine 10 generates a high bypass flow arrangement with a bypass ratio in which a large portion of the airflow which enters the fan nacelle 34 becomes bypass flow B.
- the bypass flow B communicates through the generally annular bypass flow path 40 and is discharged from the engine 10 through a variable area fan nozzle (VAFN) 42 which defines a nozzle exit area 44 between the fan nacelle 34 and the core nacelle 12 at a fan nacelle trailing edge 34 S of the fan nacelle 34 downstream of the fan section 20 .
- VAFN variable area fan nozzle
- Thrust is a function of density, velocity, and area. One or more of these parameters can be manipulated to vary the amount and direction of thrust provided by the bypass flow B.
- the VAFN 42 operates to effectively vary the area of the fan nozzle exit area 44 to selectively adjust the mass flow of the bypass flow B in response to a controller C.
- Low pressure ratio turbofans are desirable for their high propulsive efficiency. However, low pressure ratio fans may be inherently susceptible to fan stability/flutter problems at low power and low flight speeds.
- the VAFN 42 allows the engine to change to a more favorable fan operating line at low power, avoiding the instability region and still provide the relatively smaller nozzle area necessary to obtain a high-efficiency fan operating line at cruise speeds.
- the fan section 20 of the engine 10 is designed for a particular flight condition—typically cruise at 0.8 M and 35,000 feet. As the fan blades within the fan section 20 are efficiently designed at a particular fixed stagger angle for an efficient cruise condition, the VAFN 42 is operated to effectively vary the fan nozzle exit area 44 to adjust fan bypass air flow such that the angle of attack or incidence on the fan blades is maintained close to the design incidence for efficient engine operation at other flight conditions, such as landing and takeoff to thus provide optimized engine operation over a range of flight conditions with respect to performance and other operational parameters such as noise levels.
- the VAFN 42 may be separated into at least two sectors 42 A- 42 B ( FIG. 2 ) defined between the pylon P and a lower Bi-Fi splitter L which may interconnect a larger diameter fan duct reverser cowl and a smaller diameter core cowl. It should be understood that although two segments are illustrated, any number of sectors may alternatively or additionally be provided.
- the VAFN 42 selectively defines an auxiliary port system 50 with a first fan nacelle section 52 and a second fan nacelle section 54 rotationally mounted relative the first fan nacelle section 52 .
- the first fan nacelle section 52 at least partially defines the fan nacelle trailing edge 34 S with an intermittent trailing edge 60 .
- the intermittent trailing edge 60 in one disclosed non-limiting embodiment provides a saw-tooth edge which forms a multiple of ports 62 ( FIG. 3B ).
- Each of the multiple of ports 62 has a port trailing edge 62 S axially forward of the fan nacelle trailing edge 34 S.
- the second fan nacelle section 54 may include circular ring portion 64 with a multiple of doors 66 which extend therefrom and match the multiple of ports 62 .
- Each of the multiple of doors 66 has a door trailing edge 66 S aligned with the fan nacelle trailing edge 34 S ( FIG. 6A ). It should be understood that although the multiple of ports 62 and the multiple of doors 66 in the illustrated embodiment are generally rectilinear, other shapes or combinations of various shapes may alternatively be provided.
- the second fan nacelle section 54 ′ may be defined by two semi-ring portions ( FIG. 5 ) which correspond with the respective sectors 42 A- 42 B ( FIG. 2 ) defined between the pylon P and the lower Bi-Fi splitter L.
- the semi-ring portions of the second fan nacelle section 54 ′ may facilitate rotational movement thereof.
- the second fan nacelle section 54 is selectively translatable about the engine axis A relative the fixed first fan nacelle section 52 to change the effective area of the fan nozzle exit area 44 through selective opening of the ports 62 . That is, the second fan nacelle section 54 may, in one non-limiting embodiment, rotate or otherwise move about the engine axis A. As the second fan nacelle section 54 selectively translates about the engine axis A, the ports 62 are either closed by the doors 66 in the second fan nacelle section 54 ( FIG. 6A ) or are opened by offset of the second fan nacelle section 54 relative the ports 62 ( FIG. 6B ).
- each of the multiple of doors 66 match each of the multiple of ports 62 such that the fan nacelle trailing edge 34 S is continuous when the second fan nacelle section 54 is selectively translated to the closed position relative the fixed first fan nacelle section 52 ( FIG. 6A ).
- the second fan nacelle section 54 is illustrated in the disclosed non-limning embodiment as being rotatable, relative the fixed first fan nacelle section 52 , it should be understood that other translatable movement may alternatively or additionally be provided.
- the VAFN 42 communicates with the controller C to selectively translate about the engine axis A the second fan nacelle section 54 relative the first fan nacelle section 52 through an actuator system 70 to change the fan nozzle exit area 44 .
- various control systems including an engine controller or an aircraft flight control system may also be usable with the present application.
- the VAFN 42 changes the physical area and geometry of the bypass flow path 40 during particular flight modes to accommodate optimum conditions for the engine such as the Fan Pressure Ratio (FPR) that is varied in response to particular flight modes.
- FPR Fan Pressure Ratio
- the bypass flow B is effectively altered by rotation of the second fan nacelle section 54 relative the first fan nacelle section 52 between a closed position ( FIGS. 3A and 6A ) and an open position ( FIGS. 3B and 6B ).
- Rotation of the second fan nacelle section 54 to close the multiple of ports 62 of the auxiliary port system 60 decrease the fan nozzle exit area 44 toward exit area R 1 .
- Rotation of the second fan nacelle section 54 to open the ports 62 opens the auxiliary port system 60 to increase the fan nozzle exit area 44 toward exit area R 2 . That is, exit area R 2 is greater than exit area R 1 .
- an interface 68 between the second fan nacelle section 54 and the first fan nacelle section 52 includes a ring flange 72 of the circular ring portion 64 which is rotationally trapped between the first fan nacelle section 52 and a fairing 74 attached to the first fan nacelle section 52 .
- the fairing 74 defines an outer aerodynamic surface for the bypass flow B.
- the fairing 74 includes a radial faring flange 76 and the first fan nacelle section 52 includes a radial nacelle flange 78 between which the radial faring flange 76 is slidably located.
- the ring flange 72 extends in an inboard direction while the radial faring flange 76 and the radial nacelle flange 78 extend in an outboard direction relative the axis A. It should be understood that various friction reduction elements such as bearing and slider surfaces may additionally be provided.
- one non-limiting embodiment of the actuator system 70 A includes a gear rack 80 on the second fan nacelle section 54 which meshes with a pinion gear 82 .
- a slot 74 S in the faring 74 provides access to the gear rack 80 ( FIG. 10 ).
- the pinion gear 82 is translated by an actuator 84 such as a hydraulic, electric or pneumatic drive.
- the actuator 84 may alternatively be positioned within the first fan nacelle section 52 such that the actuator 84 is within the outer aerodynamic surface which bounds the bypass flow B. Rotation of the pinion gear 82 drives the gear rack 80 and thereby position the fan nacelle section 54 relative to the first fan nacelle section 52 between a closed position and an open position ( FIGS. 6A and 6B ).
- FIG. 11 another non-limiting embodiment of the actuator system 70 B includes a mount 86 on the second fan nacelle section 54 ′′ ( FIG. 12 ) which is directly actuated. That is, the mount 86 receives, for example, a push/pull input by an actuator 88 such as a linear actuator to position the fan nacelle section 54 relative to the first fan nacelle section 52 between a closed position and an open position ( FIGS. 6A and 6B ). A slot 74 S in the fairing 74 provides access for the mount 86 ( FIG. 12 ). It should be understood that although the actuator 88 is illustrated schematically and may alternatively be positioned within the first fan nacelle section 52 such that the actuator 88 is within the outer aerodynamic surface which bounds the bypass flow B.
- an actuator 88 such as a linear actuator to position the fan nacelle section 54 relative to the first fan nacelle section 52 between a closed position and an open position ( FIGS. 6A and 6B ).
- a slot 74 S in the fairing 74 provides
- another non-limiting embodiment of the actuator system 70 C includes a gear rack 90 on the second fan nacelle section 54 meshes with a pinion gear 92 driven by a gear system 94 ( FIG. 14 ).
- the gear system may include a pinion gear 96 on a common shaft 100 with the pinion gear 92 such that a worm gear 98 drives the pinion gear 96 ( FIGS. 16 and 17 ).
- the worm gear 94 is remotely driven through a flexible shaft 102 powered by an actuator 104 such as a hydraulic, electric or pneumatic drive.
- the flexible shaft 102 facilitates location of the actuator 104 in a remote location such as within the engine pylon P.
- Rotation of the flexible shaft 102 by the actuator 104 drives the worm gear 98 .
- the worm gear drives the pinion gear 96 to drive the pinion gear 92 and thus the gear rack 90 to thereby position the fan nacelle section 54 relative to the first fan nacelle section 52 between a closed position and an open position ( FIGS. 6A and 6B ).
Abstract
Description
Claims (19)
Priority Applications (2)
Application Number | Priority Date | Filing Date | Title |
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US12/909,793 US8549834B2 (en) | 2010-10-21 | 2010-10-21 | Gas turbine engine with variable area fan nozzle |
EP11186217.3A EP2444645B1 (en) | 2010-10-21 | 2011-10-21 | Gas turbine engine with variable area fan nozzle and method of operating a variable area fan nozzle |
Applications Claiming Priority (1)
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US12/909,793 US8549834B2 (en) | 2010-10-21 | 2010-10-21 | Gas turbine engine with variable area fan nozzle |
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US20120096831A1 US20120096831A1 (en) | 2012-04-26 |
US8549834B2 true US8549834B2 (en) | 2013-10-08 |
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US12/909,793 Active 2031-09-05 US8549834B2 (en) | 2010-10-21 | 2010-10-21 | Gas turbine engine with variable area fan nozzle |
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Cited By (4)
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US20180119639A1 (en) * | 2016-05-24 | 2018-05-03 | Rolls-Royce Plc | Aircraft gas turbine engine nacelle |
EP3591208A1 (en) * | 2018-07-03 | 2020-01-08 | Rolls-Royce plc | Aircraft engine fan |
US10550704B2 (en) | 2013-08-23 | 2020-02-04 | United Technologies Corporation | High performance convergent divergent nozzle |
US11440671B2 (en) * | 2019-01-24 | 2022-09-13 | Amazon Technologies, Inc. | Adjustable motor fairings for aerial vehicles |
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FR2958910B1 (en) * | 2010-04-20 | 2012-04-27 | Aircelle Sa | NACELLE FOR AIRCRAFT ENGINE WITH VARIABLE SECTION TUBE |
US8979484B2 (en) | 2012-01-05 | 2015-03-17 | Pratt & Whitney Canada Corp. | Casing for an aircraft turbofan bypass engine |
FR2996258B1 (en) * | 2012-10-01 | 2014-10-17 | Snecma | ALTERNATIVE ROTATION MIXER FOR A TURBOMACHINE CONFLUENT FLUX TUBE AND ITS STEERING PROCESS |
US9989009B2 (en) | 2012-10-31 | 2018-06-05 | The Boeing Company | Methods and apparatus for sealing variable area fan nozzles of jet engines |
US20160017815A1 (en) * | 2013-03-12 | 2016-01-21 | United Technologies Corporation | Expanding shell flow control device |
WO2016000673A1 (en) | 2014-06-30 | 2016-01-07 | Rudolf Ganz | Chevron nozzle |
US10641208B2 (en) * | 2017-11-27 | 2020-05-05 | Rohr, Inc. | Translating nozzle for mixed flow turbofan engine |
CN114087087B (en) * | 2021-10-29 | 2023-03-31 | 南京航空航天大学 | Multi-principle multi-mode pneumatic thrust vectoring nozzle and control method |
US11828248B1 (en) * | 2022-06-27 | 2023-11-28 | Pratt & Whitney Canada Corp. | Rotatably driven exhaust mixer |
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US10550704B2 (en) | 2013-08-23 | 2020-02-04 | United Technologies Corporation | High performance convergent divergent nozzle |
US20180119639A1 (en) * | 2016-05-24 | 2018-05-03 | Rolls-Royce Plc | Aircraft gas turbine engine nacelle |
US10662895B2 (en) * | 2016-05-24 | 2020-05-26 | Rolls -Royce Plc | Aircraft gas turbine engine nacelle |
EP3591208A1 (en) * | 2018-07-03 | 2020-01-08 | Rolls-Royce plc | Aircraft engine fan |
US11440671B2 (en) * | 2019-01-24 | 2022-09-13 | Amazon Technologies, Inc. | Adjustable motor fairings for aerial vehicles |
US20220371740A1 (en) * | 2019-01-24 | 2022-11-24 | Amazon Technologies, Inc. | Adjustable motor fairings for aerial vehicles |
Also Published As
Publication number | Publication date |
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EP2444645A2 (en) | 2012-04-25 |
EP2444645A3 (en) | 2013-08-21 |
US20120096831A1 (en) | 2012-04-26 |
EP2444645B1 (en) | 2016-05-04 |
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